Rotor Assembly having Collective Pitch Control

ABSTRACT

A rotor assembly for an aircraft operable to generate a variable thrust output at a constant rotational speed. The rotor assembly includes a mast rotatable at the constant speed about a mast axis. A rotor hub is coupled to and rotatable with the mast. The rotor hub includes a plurality of spindle grips extending generally radially outwardly. Each of the spindle grips is coupled to one of a plurality of rotor blades and is operable to rotate therewith about a pitch change axis. A collective pitch control mechanism is coupled to and rotatable with the rotor hub. The collective pitch control mechanism is operably associated with each spindle grip such that actuation of the collective pitch control mechanism rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades, thereby generating the variable thrust output.

TECHNICAL FIELD OF THE DISCLOSURE

The present disclosure relates, in general, to aircraft operable to transition between a forward flight mode and a vertical takeoff and landing flight mode and, in particular, to aircraft having a distributed propulsion system including a plurality of rotor assemblies, each operable to generate a variable thrust output and a variable thrust vector at a constant rotational speed.

BACKGROUND

Fixed-wing aircraft, such as airplanes, are capable of flight using wings that generate lift responsive to the forward airspeed of the aircraft, which is generated by thrust from one or more jet engines or propellers. The wings generally have an airfoil cross section that deflects air downward as the aircraft moves forward, generating the lift force to support the airplane in flight. Fixed-wing aircraft, however, typically require a runway that is hundreds or thousands of feet long for takeoff and landing.

Unlike fixed-wing aircraft, vertical takeoff and landing (VTOL) aircraft do not require runways. Instead, VTOL aircraft are capable of taking off, hovering and landing vertically. One example of VTOL aircraft is a helicopter which is a rotorcraft having one or more rotors that provide lift and thrust to the aircraft. The rotors not only enable hovering and vertical takeoff and landing, but also enable, forward, backward and lateral flight. These attributes make helicopters highly versatile for use in congested, isolated or remote areas where fixed-wing aircraft may be unable to takeoff and land. Helicopters, however, typically lack the forward airspeed of fixed-wing aircraft.

A tiltrotor aircraft is another example of a VTOL aircraft. Tiltrotor aircraft generate lift and propulsion using proprotors that are typically coupled to nacelles mounted near the ends of a fixed wing. The nacelles rotate relative to the fixed wing such that the proprotors have a generally horizontal plane of rotation for vertical takeoff, hovering and landing and a generally vertical plane of rotation for forward flight, wherein the fixed wing provides lift and the proprotors provide forward thrust. In this manner, tiltrotor aircraft combine the vertical lift capability of a helicopter with the speed and range of fixed-wing aircraft. Tiltrotor aircraft, however, typically suffer from downwash inefficiencies during vertical takeoff and landing due to interference caused by the fixed wing.

A further example of a VTOL aircraft is a tiltwing aircraft that features a rotatable wing that is generally horizontal for forward flight and rotates to a generally vertical orientation for vertical takeoff and landing. Propellers are coupled to the rotating wing to provide the required vertical thrust for takeoff and landing and the required forward thrust to generate lift from the wing during forward flight. The tiltwing design enables the slipstream from the propellers to strike the wing on its smallest dimension, thus improving vertical thrust efficiency as compared to tiltrotor aircraft. Tiltwing aircraft, however, are more difficult to control during hover as the vertically tilted wing provides a large surface area for crosswinds typically requiring tiltwing aircraft to have either cyclic rotor control or an additional thrust station to generate a moment.

SUMMARY

In a first aspect, the present disclosure is directed to a rotor assembly for an aircraft operable to generate a variable thrust output at a constant rotational speed. The rotor assembly includes a mast that is rotatable at the constant speed about a mast axis. A rotor hub is coupled to and rotatable with the mast. The rotor hub includes a plurality of spindle grips that extend generally radially outwardly. Each of the spindle grips is operable to rotate about a pitch change axis. Each of a plurality of rotor blades is coupled to one of the spindle grips and is rotatable therewith about the respective pitch change axis. A collective pitch control mechanism is coupled to the rotor hub and is rotatable therewith. The collective pitch control mechanism is operably associated with each of the spindle grips. In operation, actuation of the collective pitch control mechanism rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades, thereby generating the variable thrust output.

In certain embodiments, the collective pitch control mechanism may include a spider assembly having a plurality of arms each of which mate with one of the spindle grips such that translation of the spider assembly relative to the rotor hub rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades. In such embodiments, the collective pitch control mechanism may include at least one actuator operable to translation the spider assembly relative to the rotor hub. Also, in such embodiments, each spindle grip may include a pinion gear and each arm may include a rack gear that is operable to mate with one of the pinion gears. In some embodiments, the collective pitch control mechanism may include a ring assembly that mates with each of the spindle grips such that rotation of the ring assembly relative to the rotor hub rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades. In such embodiments, the ring assembly may include an input gear and the collective pitch control mechanism may include at least one actuator having a drive gear that is operable to mate with the input gear and rotate the ring assembly relative to the rotor hub. Also, in such embodiments, each spindle grip may include a pinion gear and the ring assembly may include an output gear that is operable to mate with each of the pinion gears.

In certain embodiments, the rotor assembly may include a joint assembly, such as a constant velocity joint assembly, that provides a torque path from the mast to the rotor hub. In such embodiments, the rotor hub may include a drive arm assembly that is operable to rotate with and translate relative to the constant velocity joint assembly during rotary operations. In some embodiments, the rotor assembly may include a ball joint positioned about and non rotatable with the mast and a tilt control assembly positioned on and having a tilting degree of freedom relative to the ball joint wherein the tilt control assembly is non rotatable with the mast. In such embodiments, the rotor hub may be rotatably coupled to and tiltable with the tilt control assembly such that actuation of the tilt control assembly changes the rotational plane of the rotor hub relative to the mast axis.

In a second aspect, the present disclosure is directed to an aircraft that includes an airframe, a propulsion system having a plurality of propulsion assemblies each with a rotor assembly and a flight control system operable to independently control each of the propulsion assemblies. Each of the rotor assemblies includes a mast that rotatable relative to the airframe about a mast axis. A rotor hub is coupled to and rotatable with the mast. The rotor hub includes a plurality of spindle grips that extend generally radially outwardly. Each of the spindle grips is operable to rotate about a pitch change axis. Each of a plurality of rotor blades is coupled to one of the spindle grips and is rotatable therewith about the respective pitch change axis. A collective pitch control mechanism is coupled to the rotor hub and is rotatable therewith. The collective pitch control mechanism is operably associated with each of the spindle grips. In operation, actuation of the collective pitch control mechanism rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades, thereby generating the variable thrust output.

In some embodiments, the airframe may include first and second wings each having an M-wing design with two leading apexes such that each of the rotor assemblies is positioned proximate one of the leading apexes. In certain embodiments, a passenger pod assembly may be rotatably attachable to the airframe such that the passenger pod assembly remains in a generally horizontal attitude in a vertical takeoff and landing fight mode, a forward flight mode and transitions therebetween. In some embodiments, the flight control system may command operation of the propulsion assemblies responsive to at least one of onboard pilot flight control, remote flight control, autonomous flight control and combinations thereof.

BRIEF DESCRIPTION OF THE DRAWINGS

For a more complete understanding of the features and advantages of the present disclosure, reference is now made to the detailed description along with the accompanying figures in which corresponding numerals in the different figures refer to corresponding parts and in which:

FIGS. 1A-1F are schematic illustrations of an aircraft in accordance with embodiments of the present disclosure;

FIGS. 2A-2B are block diagrams of propulsion assemblies for an aircraft in accordance with embodiments of the present disclosure;

FIGS. 3A-3I are schematic illustrations of an aircraft in a sequential flight operating scenario in accordance with embodiments of the present disclosure;

FIGS. 4A-4D are schematic illustrations of a propulsion assembly for an aircraft in accordance with embodiments of the present disclosure;

FIGS. 5A-5H are schematic illustrations of an aircraft in various ground maneuver configurations in accordance with embodiments of the present disclosure;

FIGS. 6A-6D are schematic illustrations of a propulsion assembly for an aircraft depicting the tilting degree of freedom and thrust vector generation of a rotor assembly in accordance with embodiments of the present disclosure;

FIGS. 7A-7D are various views depicting the connections between propulsion assemblies and an airframe of an aircraft in accordance with embodiments of the present disclosure;

FIG. 8 is a block diagram of an aircraft control system in accordance with embodiments of the present disclosure;

FIGS. 9A-9B are isometric and exploded views of a rotor assembly operable to generate a variable thrust output and a variable thrust vector at a constant rotational speed for an aircraft in accordance with embodiments of the present disclosure; and

FIGS. 10A-10B are isometric and exploded views of a rotor assembly operable to generate a variable thrust output and a variable thrust vector at a constant rotational speed for an aircraft in accordance with embodiments of the present disclosure.

DETAILED DESCRIPTION

While the making and using of various embodiments of the present disclosure are discussed in detail below, it should be appreciated that the present disclosure provides many applicable inventive concepts, which can be embodied in a wide variety of specific contexts. The specific embodiments discussed herein are merely illustrative and do not delimit the scope of the present disclosure. In the interest of clarity, not all features of an actual implementation may be described in the present disclosure. It will of course be appreciated that in the development of any such actual embodiment, numerous implementation-specific decisions must be made to achieve the developer's specific goals, such as compliance with system-related and business-related constraints, which will vary from one implementation to another. Moreover, it will be appreciated that such a development effort might be complex and time-consuming but would be a routine undertaking for those of ordinary skill in the art having the benefit of this disclosure.

In the specification, reference may be made to the spatial relationships between various components and to the spatial orientation of various aspects of components as the devices are depicted in the attached drawings. However, as will be recognized by those skilled in the art after a complete reading of the present disclosure, the devices, members, apparatuses, and the like described herein may be positioned in any desired orientation. Thus, the use of terms such as “above,” “below,” “upper,” “lower” or other like terms to describe a spatial relationship between various components or to describe the spatial orientation of aspects of such components should be understood to describe a relative relationship between the components or a spatial orientation of aspects of such components, respectively, as the device described herein may be oriented in any desired direction.

Referring to FIGS. 1A-1F in the drawings, various views of an aircraft 10 having a versatile propulsion system are depicted. In the illustrated embodiment, aircraft 10 including an airframe 12 having wings 14, 16 each have an airfoil cross-section that generates lift responsive to the forward airspeed of aircraft 10. Wings 14, 16 may be formed as single members or may be formed from multiple wing sections. The outer skins for wings 14, 16 are preferably formed from high strength and lightweight materials such as fiberglass fabric, carbon fabric, fiberglass tape, carbon tape and combinations thereof that may be formed by curing together a plurality of material layers.

Extending generally perpendicularly between wings 14, 16 are truss structures depicted as pylons 18, 20. Pylons 18, 20 are preferably formed from high strength and lightweight materials such as fiberglass fabric, carbon fabric, fiberglass tape, carbon tape and combinations thereof that may be formed by curing together a plurality of material layers. Preferably, wings 14, 16 and pylons 18, 20 are securably attached together at the respective intersections by bolting, bonding and/or other suitable technique such that airframe 12 becomes a unitary member. As illustrated, wings 14, 16 are polyhedral wings with wing 14 having anhedral sections 14 a, 14 b and with wing 16 having dihedral sections 16 a, 16 b, as best seen in FIG. 1B. In this design, any liquid fuel stored in the anhedral sections 14 a, 14 b or dihedral sections 16 a, 16 b of wings 14, 16 will gravity feed to sump locations. Wings 14, 16 preferably include central passageways operable to contain energy sources and communication lines. For example, as best seen in FIG. 1A, wing 14 includes energy sources 22 a, 22 b and communication lines 24 a, 24 b. Energy sources 22 a, 22 b may be liquid fuel, batteries or other suitable energy sources. In the case of liquid fuel energy sources 22 a, 22 b, communication lines 24 a, 24 b form a fluid distribution network wherein communication lines 24 a, 24 b access energy sources 22 a, 22 b proximate the outboard end of wing 14 due to the gravity feed enabled by anhedral sections 14 a, 14 b. A pumping system may be used to move the liquid fuel from the gravity sumps to the desired use location. In the case of liquid fuel energy sources in wing 16, the communication lines in wing 16 would access the liquid fuel energy sources proximate the inboard end of wing 16 due to the gravity feed enabled by dihedral sections 16 a, 16 b.

In the illustrated embodiment, the versatile propulsion system includes a plurality of interchangeably propulsion assemblies 26 a, 26 b, 26 c, 26 d that are independently attachable to and detachable from airframe 12. As illustrated, propulsion assemblies 26 a, 26 b, 26 c, 26 d are positioned on airframe 12 in a close coupled configuration. As illustrated, the versatile propulsion system includes four independently operating propulsion assemblies 26 a, 26 b, 26 c, 26 d. It should be noted, however, that a versatile propulsion system of the present disclosure could have any number of independent propulsion assemblies including six, eight, twelve, sixteen or other number of independent propulsion assemblies. In the illustrated embodiment, propulsion assemblies 26 a, 26 b are securably attached to airframe 12 in a high wing configuration and propulsion assemblies 26 c, 26 d are securably attached to airframe 12 in a low wing configuration by bolting or other suitable technique, as best seen in FIG. 1B. Preferably, each propulsion assembly 26 a, 26 b, 26 c, 26 d includes a nacelle 28 a, 28 b, 28 c, 28 d, as best seen in FIG. 1F, that houses a power source, an engine or motor, a drive system, a rotor hub, actuators and an electronics node including, for example, controllers, sensors and communications elements as well as other components suitable for use in the operation of a propulsion assembly. Each propulsion assembly 26 a, 26 b, 26 c, 26 d also includes a tail assembly 46 a, 46 b, 46 c, 46 d having an active aerosurface 48 a, 48 b, 48 c, 48 d, as best seen in FIGS. 1E and 1F. In addition, each propulsion assembly 26 a, 26 b, 26 c, 26 d has a rotor assembly including the rotor hub having a plurality of grips such as spindle grips and a proprotor 38 a, 38 b, 38 c, 38 d depicted as having three rotor blades each of which is coupled to one of the spindle grips of the respective rotor hub such that the rotor blades are operable to rotate with the spindle grips about respective pitch change axes, as discussed herein.

Aircraft 10 has a liquid fuel flight mode, wherein energy is provided to each of the propulsion assemblies from liquid fuel. For example, in this configuration, each of the propulsion assemblies may be represented by propulsion assembly 26 a of FIGS. 1A and 2A. As illustrated, propulsion assembly 26 a includes a nacelle 28 a, one or more fuel tanks 30 a, an internal combustion engine 32 a, a drive system 34 a, a rotor hub 36 a, a proprotor 38 a and an electronics node 40 a. In the liquid fuel flight mode, the fuel tanks of the propulsion assemblies may be connected to the fluid distribution network of the airframe and serve as feeder tanks for the IC engines. Alternatively, the liquid fuel system may be a distributed system wherein liquid fuel for each propulsion assembly is fully self-contained within the fuel tanks positioned within the nacelles, in which case, the wet wing system described above may not be required. The IC engines may be powered by gasoline, jet fuel, diesel or other suitable liquid fuel. The IC engines may be rotary engines such as dual rotor or tri rotor engines or other high power-to-weight ratio engines. The drive systems may include multistage transmissions operable for reduction drive such that optimum engine rotation speed and optimum proprotor rotation speed are enabled. The drive systems may utilize high-grade roller chains, spur and bevel gears, v-belts, high strength synchronous belts or the like. As one example, the drive system may be a two-stage cogged belt reducing transmission including a 3 to 1 reduction in combination with a 2 to 1 reduction resulting in a 6 to 1 reduction between the engine and the rotor hub.

Aircraft 10 also has an electric flight mode, wherein energy is provided to each of the propulsion assemblies from an electric power source. For example, in this configuration, each of the propulsion assemblies may be represented by propulsion assembly 26 b of FIGS. 1A and 2B. As illustrated, propulsion assembly 26 b includes a nacelle 28 b, one or more batteries 30 b, an electric motor 32 b, a drive system 34 b, a rotor hub 36 b, a proprotor 38 b and an electronics node 40 b. In the electric flight mode, the electric motors of each propulsion assembly are preferably operated responsive to electrical energy from the battery or batteries disposed with that nacelle, thereby forming a distributed electrical system. Alternatively or additionally, electrical power may be supplied to the electric motors and/or the batteries disposed with the nacelles from the energy sources, such as energy sources 22 a, 22 b, carried by airframe 12 via the communication lines, such as communication lines 24 a, 24 b.

Aircraft 10 also has a mixed flight mode, wherein energy is provided to some of the propulsion assemblies from an electric power source and energy is provided to other of the propulsion assemblies from liquid fuel. The mixed flight mode of aircraft 10 is evident from the illustrated embodiment of FIG. 1A. As another alternative, some or all of the engines of the propulsion assembly 26 a, 26 b, 26 c, 26 d may be hydraulic motors operated responsive to a distributed hydraulic fluid system wherein high pressure hydraulic sources or generators are housed within each nacelle. Alternatively or additionally, a common hydraulic fluid system integral to or carried by airframe 12 may be used.

To transition aircraft 10 among liquid fuel flight mode, electric flight mode and mixed flight mode, propulsion assembly 26 a, 26 b, 26 c, 26 d are preferably standardized and interchangeable units that are most preferably line replaceable units enabling easy installation and removal from airframe 12, as discussed herein. In addition, regardless of the current flight mode configuration of aircraft 10, the use of line replaceable units is beneficial in maintenance situations if a fault is discovered with one of the propulsion assemblies. In this case, the faulty propulsion assembly can be decoupled from airframe 12 by simple operations such as unbolting structural members, disconnecting power, data and/or fluid couplings and other suitable procedures. Another propulsion assembly can then be attached to airframe 12 by connecting power, data and/or fluid couplings, bolting structural members together and other suitable procedures.

The rotor assemblies of each propulsion assembly 26 a, 26 b, 26 c, 26 d are preferably lightweight, rigid members that may include swashyoke mechanisms operable for collective pitch control and thrust vectoring. Proprotors 38 a, 38 b, 38 c, 38 d each include a plurality of proprotor blades that are securably attached to spindle grips of the respective rotor hub. The blades are preferably operable for collective pitch control and may additional be operable for cyclic pitch control. As an alternative, the pitch of the blades may be fixed, in which case, thrust is determined by changes in the rotational velocity of the proprotors. In the illustrated embodiment, the rotor hubs have a tilting degree of freedom to enable thrust vectoring.

To accommodate the tilting degree of freedom of the rotor hubs, wings 14, 16 have a unique swept wing design, which is referred to herein as an M-wing design. For example, as best seen in FIG. 1A, wing 14 has swept forward portions 14 c, 14 d and swept back portions 14 e, 14 f. Propulsion assembly 26 a is coupled to a wing stanchion positioned between swept forward portion 14 c and swept back portion 14 e. Likewise, propulsion assembly 26 b is coupled to a wing stanchion positioned between swept forward portion 14 d and swept back portion 14 f. Wing 16 has a similar M-wing design with propulsion assemblies 26 c, 26 d similarly coupled to wing stanchions positioned between swept forward and swept back portions. In this configuration, each rotor hub is operable to pivot about a mast axis, such as mast axis 42 a and mast axis 42 b, to control the direction of the thrust vector while avoiding any interference between proprotor 38 a, 38 b, 38 c, 38 d and wings 14, 16. In the illustrated embodiment, the maximum angle of the thrust vector may preferably be between about 10 degrees and about 30 degrees, may more preferably be between about 15 degrees and about 25 degrees and may most preferably be about 20 degrees. In embodiments having a maximum thrust vector angle of 20 degrees, the thrust vector may be resolved to any position within a 20-degree cone swung about the mast centerline axis. Notably, using a 20-degree thrust vector yields a lateral component of thrust that is about 34 percent of total thrust. Even though the propulsion assemblies of the present disclosure have been described as having certain nacelles, power sources, engines, drive systems, rotor hubs, proprotors and tail assemblies, it is to be understood by those having ordinary skill in the art that propulsion assemblies having other components or combinations of components suitable for use in a versatile propulsion system are also possible and are considered to be within the scope of the present disclosure.

As best seen in FIGS. 1A, 1C and 1E, aircraft 10 includes landing gear depicted as including wheels 44 a, 44 b, 44 c, 44 d. The landing gear may be passively operated pneumatic landing struts or actively operated telescoping landing struts disposed within tail assemblies 46 a, 46 b, 46 c, 46 d of propulsion assemblies 26 a, 26 b, 26 c, 26 d. As discussed herein, wheels 44 a, 44 b, 44 c, 44 d enable aircraft 10 to taxi and perform other ground maneuvers. The landing gear may provide a passive brake system or may include active brakes such as an electromechanical braking system or a manual braking system to facilitate parking as required during ground operations and/or passenger ingress and egress. In the illustrated embodiment, each tail assembly 46 a, 46 b, 46 c, 46 d includes an active aerosurface 48 a, 48 b, 48 c, 48 d that is controlled by an active aerosurface control module of a flight control system 40. During various flight operations, active aerosurfaces 48 a, 48 b, 48 c, 48 d may operate as vertical stabilizers, horizontal stabilizers, rudders and/or elevators to selectively provide pitch control and yaw control to aircraft 10.

As best seen in FIGS. 1A and 2A-2B, each of the propulsion assemblies 26 a, 26 b, 26 c, 26 d includes one or more actuators, such as actuators 52 a disposed within nacelle 28 a and actuators 52 b disposed within nacelle 28 b, that are operable to change the orientation, the configuration and/or the position of the active aerosurface 48 a, 48 b, 48 c, 48 d. For each propulsion assembly and responsive to commands from the active aerosurface control module of flight control system 40, a first actuator is operable to rotate the tail assembly relative to the nacelle (see FIG. 4A). This operation also rotates the active aerosurface relative to the nacelle such that the active aerosurfaces 48 a, 48 b, 48 c, 48 d may be generally parallel to wings 14, 16, as best seen in FIG. 1E, the active aerosurfaces 48 a, 48 b, 48 c, 48 d may be generally perpendicular to wings 14, 16, as best seen in FIG. 1F, or the active aerosurfaces 48 a, 48 b, 48 c, 48 d may be located in any orientation therebetween. Active aerosurfaces 48 a, 48 b, 48 c, 48 d may be fixed, in which case, when active aerosurfaces 48 a, 48 b, 48 c, 48 d are generally perpendicular to wings 14, 16 they may be referred to as vertical stabilizers that provide yaw control to aircraft 10. Similarly, when active aerosurfaces 48 a, 48 b, 48 c, 48 d are generally parallel to wings 14, 16 they may be referred to as horizontal stabilizers that provide pitch control to aircraft 10.

For each propulsion assembly and responsive to commands from the active aerosurface control system of flight control module 40, a second actuator may be operable to tilt all or a portion of each active aerosurface relative to the tail assembly (see FIGS. 4C-4D). In this case, when active aerosurfaces 48 a, 48 b, 48 c, 48 d are generally perpendicular to wings 14, 16 and are tilted relative to tail assemblies 46 a, 46 b, 46 c, 46 d, active aerosurfaces 48 a, 48 b, 48 c, 48 d may be referred to as rudders that provide yaw control to aircraft 10. Likewise, when active aerosurfaces 48 a, 48 b, 48 c, 48 d are generally parallel to wings 14, 16 and are tilted relative to tail assemblies 46 a, 46 b, 46 c, 46 d, active aerosurfaces 48 a, 48 b, 48 c, 48 d may be referred to as elevators that provide pitch control to aircraft 10.

For each propulsion assembly and responsive to commands from the active aerosurface control module of flight control system 40, a third actuator is operable to translate the tail assembly relative to the nacelle between a retracted configuration and an extended configuration (see FIG. 4B). In the extended configuration of tail assemblies 46 a, 46 b, 46 c, 46 d relative to nacelles 28 a, 28 b, 28 c, 28 d, as best seen in FIG. 1F, active aerosurfaces 48 a, 48 b, 48 c, 48 d are able to generate a greater moment, which can be beneficial in stabilizing aircraft 10. It should be noted that during certain flight maneuvers, it may be beneficial to have active aerosurfaces 48 a, 48 b, 48 c, 48 d operating as vertical stabilizers and/or rudders such as during forward flight maneuvers, as best seen in FIG. 1F. Likewise, during other flight maneuvers, it may be beneficial to have active aerosurfaces 48 a, 48 b, 48 c, 48 d operating as horizontal stabilizers and/or elevators such as during hover flight maneuvers, vertical takeoff and land flight maneuvers and transition between forward flight and VTOL flight, as best seen in FIG. 1E. In addition, it may be beneficial to have some of active aerosurfaces 48 a, 48 b, 48 c, 48 d operating as horizontal stabilizers and/or elevators and other of active aerosurfaces 48 a, 48 b, 48 c, 48 d operating as vertical stabilizers and/or rudders during certain flight maneuvers. Further, it may be beneficial to have one or more of active aerosurfaces 48 a, 48 b, 48 c, 48 d operating in a position between the horizontal stabilizer position and the vertical stabilizer position during certain flight maneuvers.

Flight control system 40 of aircraft 10, such as a digital flight control system, may be located within a central passageway of wing 14, as best seen in FIG. 1A. In the illustrated embodiment, flight control system 40 is a triply redundant flight control system including three independent flight control computers. Use of triply redundant flight control system 40 having redundant components improves the overall safety and reliability of aircraft 10 in the event of a failure in flight control system 40. Flight control system 40 preferably includes non-transitory computer readable storage media including a set of computer instructions executable by one or more processors for controlling the operation of the versatile propulsion system. Flight control system 40 may be implemented on one or more general-purpose computers, special purpose computers or other machines with memory and processing capability. For example, flight control system 40 may include one or more memory storage modules including, but is not limited to, internal storage memory such as random access memory, non-volatile memory such as read only memory, removable memory such as magnetic storage memory, optical storage, solid-state storage memory or other suitable memory storage entity. Flight control system 40 may be a microprocessor-based system operable to execute program code in the form of machine-executable instructions. In addition, flight control system 40 may be selectively connectable to other computer systems via a proprietary encrypted network, a public encrypted network, the Internet or other suitable communication network that may include both wired and wireless connections.

Flight control system 40 communicates via a communications network 54 with the electronics nodes of each propulsion assembly 26 a, 26 b, 26 c, 26 d, such as electronics node 40 a of propulsion assembly 26 a and electronics node 40 b of propulsion assembly 26 b, as best seen in FIGS. 1A and 2A-2B. Flight control system 40 receives sensor data from and sends flight command information to the electronics nodes of each propulsion assembly 26 a, 26 b, 26 c, 26 d such that each propulsion assembly 26 a, 26 b, 26 c, 26 d may be individually and independently controlled and operated. In both manned and unmanned missions, flight control system 40 may autonomously control some or all aspects of flight operation for aircraft 10. Flight control system 40 is also operable to communicate with remote systems, such as a transportation services provider system via a wireless communications protocol. The remote system may be operable to receive flight data from and provide commands to flight control system 40 to enable remote flight control over some or all aspects of flight operation for aircraft 10, in both manned and unmanned missions.

Aircraft 10 includes a pod assembly, illustrated as passenger pod assembly 50, that is selectively attachable to airframe 12 between pylons 18, 20. In the illustrated embodiment, pylons 18, 20 include receiving assemblies for coupling with pod assembly 50. Preferably, the connection between pylons 18, 20 and pod assembly 50 allows pod assembly 50 to rotate and translate relative to airframe 12 during flight operations. In addition, one or more communication channels may be established between pod assembly 50 and airframe 12 when pod assembly 50 is attached therewith. For example, a quick disconnect harness may be coupled between pod assembly 50 and airframe 12 to allow a pilot within pod assembly 50 to receive flight data from and provide commands to flight control system 40 to enable onboard pilot control over some or all aspects of flight operation for aircraft 10.

As best seen in FIG. 1E, aircraft 10 is in a vertical takeoff and landing mode. As illustrated, wings 14, 16 are generally above pod assembly 50 with wing 14 forward of and wing 16 aft of pod assembly 50 and with wings 14, 16 disposed in generally the same horizontal plane. As noted, flight control system 40 independently controls and operates each propulsion assembly 26 a, 26 b, 26 c, 26 d. In one example, flight control system 40 is operable to independently control collective pitch and adjust the thrust vector of each propulsion assembly 26 a, 26 b, 26 c, 26 d, which can be beneficial in stabilizing aircraft 10 during vertical takeoff, vertical landing and hover. As best seen in FIG. 1F, aircraft 10 is in a forward flight mode. Wings 14, 16 are generally forward of pod assembly 50 with wing 14 below and wing 16 above pod assembly 50 and with wings 14, 16 disposed in generally the same vertical plane. In the illustrated embodiment, the proprotor blades of propulsion assemblies 26 a, 26 d rotate counterclockwise while the proprotor blades of propulsion assemblies 26 b, 26 c rotate clockwise to balance the torque of aircraft 10.

Referring next to FIGS. 3A-3I in the drawings, a sequential flight-operating scenario of aircraft 10 is depicted. In the illustrated embodiment, passenger pod assembly 50 is attached to airframe 12. It is noted, however, that passenger pod assembly 50 may be selectively disconnected from airframe 12 such that a single airframe can be operably coupled to and decoupled from numerous passenger pod assemblies for numerous missions over time. As best seen in FIG. 3A, aircraft 10 is positioned on the ground with the tail assemblies 46 positioned in a passive brake configuration, as discussed herein. When aircraft 10 is ready for a mission, flight control system 40 commences operations to provide flight control to aircraft 10 which may be autonomous flight control, remote flight control, onboard pilot flight control or any combination thereof. For example, it may be desirable to utilize onboard pilot flight control during certain maneuvers such as takeoff and landing but rely on remote or autonomous flight control during hover, forward flight and/or transitions between forward flight and VTOL operations. As best seen in FIG. 3B, aircraft 10 is in its vertical takeoff and landing mode and has lifted pod assembly 50 into the air. Preferably, each tail assembly 46 is rotated about the mast axis 56 of the respective nacelle 28 (see FIG. 4A) such that active aerosurfaces 48 are generally parallel with wings 14, 16 to aid in stabilization during hover and to be properly positioned to provide pitch control during the transition to forward flight. After vertical assent to the desired elevation, aircraft 10 may begin the transition from vertical takeoff to forward flight.

As best seen in FIGS. 3B-3D, as aircraft 10 transitions from vertical takeoff and landing mode to forward flight mode, airframe 12 rotates about pod assembly 50 such that pod assembly 50 is maintained in a generally horizontal attitude for the safety and comfort of passengers, crew and/or cargo carried in pod assembly 50. This is enabled by a passive and/or active connection between airframe 12 and pod assembly 50. For example, a gimbal assembly may be utilized to allow passive orientation of pod assembly 50 relative to airframe 12. This may be achieved due to the shape and the center of gravity of pod assembly 50 wherein aerodynamic forces and gravity tend to bias pod assembly 50 toward the generally horizontal attitude. Alternatively or additionally, a gear assembly, a clutch assembly or other suitably controllable rotating assembly may be utilized that allows for pilot controlled, remote controlled or autonomously controlled rotation of pod assembly 50 relative to airframe 12 as aircraft 10 transitions from vertical takeoff and landing mode to forward flight mode. Preferably, as best seen in FIG. 3D, each tail assembly 46 is rotated about the mast axis 56 of the respective nacelle 28 (see FIG. 4A) such that active aerosurfaces 48 are generally perpendicular with wings 14, 16 to aid in stabilization and provide yaw control during forward flight.

As best seen in FIGS. 3D-3E, once aircraft 10 has completed the transition to forward flight mode, it may be desirable to adjust the center of gravity of aircraft 10 to improve its stability and efficiency. In the illustrated embodiment, this can be achieved by shifting pod assembly 50 forward relative to airframe 12 using an active connection between airframe 12 and pod assembly 50. For example, rotation of a gear assembly of pod assembly 50 relative to a rack assembly of airframe 12 or other suitable translation system may be used to shift pod assembly 50 forward relative to airframe 12 under pilot control, remote control or autonomous control. Preferably, as best seen in FIG. 3E, each tail assembly 46 is extended relative to the respective nacelle 28 (see FIG. 4B) such that active aerosurfaces 48 provide greater stabilization and yaw control during forward flight.

When aircraft 10 begins its approaches to the destination, pod assembly 50 is preferably returned to the aft position relative to airframe 12 and each tail assembly 46 is preferably retracted (see FIG. 4B) and rotated (see FIG. 4A) relative to the respective nacelle 28 such that active aerosurfaces 48 are generally parallel with wings 14, 16, as best seen in FIG. 3F. Aircraft 10 may now begin its transition from forward flight mode to vertical takeoff and landing mode. As best seen in FIGS. 3F-3H, during the transition from forward flight mode to vertical takeoff and landing flight mode, airframe 12 rotates about pod assembly 50 such that pod assembly 50 is maintained in the generally horizontal attitude for the safety and comfort of passengers, crew and/or cargo carried in pod assembly 50. Once aircraft 10 has completed the transition to vertical takeoff and landing flight mode, as best seen in FIG. 3H, aircraft 10 may commence its vertical descent to a surface. After landing, aircraft 10 may engage in ground maneuvers, if desired. Upon completion of any ground maneuvers, each tail assembly 46 is rotated about the mast axis 56 of the respective nacelle 28 (see FIG. 4A) to return to the passive brake configuration, as best seen in FIG. 3I.

Referring next to FIGS. 5A-5H and 6A-6D, the omnidirectional ground maneuver capabilities of aircraft 10 will now be described. As illustrated, aircraft 10 includes an airframe 12 having wings 14, 16. A plurality of propulsion assemblies, referred to individually and collectively as propulsion assembly 26 or propulsion assemblies 26, are attached to airframe 12. Two of the propulsion assemblies 26 are coupled to wing 14 and two of the propulsion assemblies 26 are coupled to wing 16. In the illustrated embodiment, propulsion assemblies 26 are generally symmetrically positioned or disposed about an aircraft rotational axis 58 to provide stability to aircraft 10. Each of the propulsion assemblies 26 includes a nacelle 28 having a mast axis 56, a rotor assembly 60 having a tilting degree of freedom relative to mast axis 56 and a tail assembly 46 rotatable about mast axis 56. It should be noted that each rotor assembly 60 includes a rotor hub 36 and a proprotor 38, as best see in FIG. 1A. Each tail assembly 46 forms a landing gear including at least one wheel 44 having a rotational axis 62. As discussed herein, a flight control system 40 is operable to independently control each of the propulsion assemblies 26 including tilting each rotor assembly 60 and rotating each tail assembly 46. For each propulsion assembly 20, flight control system 40 is operable to tilt rotor assembly 60 in any direction relative to mast axis 56, as best seen in FIGS. 6A-6D, enabling a thrust vector 64 to be resolved within a thrust vector cone relative to mast axis 56. In the illustrated embodiment, thrust vector 64 has a maximum angle relative to mast axis 56 of about twenty degrees. When propulsion assemblies 26 are being operated including rotation of rotor assemblies 60 about mast axis 56 and tilting of rotor assemblies 60 relative to mast axis 56, the thrust vectors 64 generated by rotor assemblies 60 have a vertical component 66 and a horizontal component 68. Importantly, the horizontal component 68 of thrust vectors 64 is operable to provide the energy source required to roll aircraft 10 during ground maneuvers.

The omnidirectional ground maneuver capabilities of aircraft 10 are achieved by controlling thrust vectors 66 of rotor assemblies 60 relative to rotational axes 62 of wheels 44 of tail assemblies 46. For example, for each propulsion assembly 26, rotor assembly 60 and tail assembly 46 have complementary configurations in which the horizontal component 68 of thrust vector 64 is generally perpendicular to rotational axis 62 of wheel 44. In this case, the thrust is pushing aircraft 10 in a direction that efficiently turns wheel 44. In addition, for each propulsion assembly 26, rotor assembly 60 and tail assembly 46 have non complementary configurations in which the horizontal component 68 of thrust vector 64 is not perpendicular to rotational axis 62 of wheel 44. In this case, the thrust may be pushing aircraft 10 in a direction that inefficiently turns wheel 44 or in a direction that does not turn wheel 44.

In FIG. 5A, rotor assemblies 60 and tail assemblies 46 have complementary configurations in which the horizontal components 68 of thrust vectors 64 are generally perpendicular to the respective rotational axes 62 of wheels 44. Rotational axes 62 of wheels 44 are generally parallel with wings 14, 16 which creates a fore/aft ground maneuver configuration, as indicated by the directional arrow intersecting aircraft rotational axis 58. In the illustrated embodiment, aircraft 10 would move in the forward direction due to the aftward direction of the horizontal components 68 of thrust vectors 64. As discussed herein, wings 14, 16 are polyhedral wings having anhedral and dihedral sections. Thus, the use of terms such as “generally parallel,” “generally perpendicular” and “generally congruent” take into account such angular variations.

In FIG. 5B, rotor assemblies 60 and tail assemblies 46 have complementary configurations. Rotational axes 62 of wheels 44 are generally perpendicular with wings 14, 16 which creates a lateral ground maneuver configuration, as indicated by the directional arrow intersecting aircraft rotational axis 58. In the illustrated embodiment, aircraft 10 would move to its left, when viewed from the front, due to the horizontal components 68 of thrust vectors 64 pointing to the right. FIGS. 5C and 5D illustrate tail assemblies 46 in two of a plurality of radial ground maneuver configurations, as indicated by the directional arrows intersecting respective aircraft rotational axes 58. Rotor assemblies 60 and tail assemblies 46 have complementary configurations and each of the rotational axes 62 of wheels 44 has been rotated to the same relative orientation between generally parallel and generally perpendicular with wings 14, 16. In this manner, aircraft 10 is operable for omnidirectional ground maneuvers.

FIG. 5E illustrates tail assemblies 46 in a turning ground maneuver configuration, as indicated by the directional arrow intersecting aircraft rotational axis 58. Aircraft 10 is being urged to the left, when viewed from the front, due to the horizontal components 68 of thrust vectors 64 pointing to the right. As with all the previous examples, rotor assemblies 60 have a common tilting configuration in which each of the rotor assemblies 60 is tilting in the same direction. In this case, however, some of the rotor assemblies 60 and tail assemblies 46 have complementary configurations while others of the rotor assemblies 60 and tail assemblies 46 have non complementary configurations. More specifically, as aircraft 10 moves to the left, the rear rotor assemblies 60 and tail assemblies 46 have complementary configurations in which rotational axes 62 of wheels 44 are generally perpendicular with wings 14, 16. The front rotor assemblies 60 and tail assemblies 46 have non complementary configurations in which rotational axes 62 of wheels 44 have been rotated to a configuration between generally parallel and generally perpendicular with wings 14, 16, thus providing front wheel steering for aircraft 10.

FIG. 5F illustrates tail assemblies 46 in an alternate turning ground maneuver configuration, as indicated by the directional arrow intersecting aircraft rotational axis 58. Aircraft 10 is being urged forward due to the horizontal components 68 of thrust vectors 64 pointing aftward. Rotor assemblies 60 have a common tilting configuration in which each of the rotor assemblies 60 is tilting in the same direction. The front rotor assemblies 60 and tail assemblies 46 have complementary configurations in which rotational axes 62 of wheels 44 are generally parallel with wings 14, 16. The rear rotor assemblies 60 and tail assemblies 46 have non complementary configurations in which rotational axes 62 of wheels 44 have been rotated to a configuration between generally parallel and generally perpendicular with wings 14, 16, thus providing rear wheel steering for aircraft 10. Alternatively or additionally, tail assemblies 46 may have a four wheel steering ground maneuver configuration.

FIG. 5G illustrates tail assemblies 46 in a rotation ground maneuver configuration, as indicated by the directional arrow around aircraft rotational axis 58. Rotor assemblies 60 and tail assemblies 46 have complementary configurations. Rotor assemblies 60, however, have a non common tilting configuration in which rotor assemblies 60 are titling in a plurality of directions and, in the illustrated embodiment, are each tilting in a different direction. In addition, rotational axis 62 of each wheel 44 intersects rotational axis 58 of aircraft 10. In this rotation ground maneuver configuration of rotor assemblies 60 and tail assemblies 46, aircraft 10 will rotate about rotational axis 58 in a counter clockwise direction, when viewed from above.

FIG. 5H illustrates tail assemblies 46 in a passive brake configuration, as indicated by the stop symbol intersecting aircraft rotational axis 58. Rotational axes 62 of adjacent wheels 44 are generally perpendicular with each other. In addition, rotational axis 62 of each wheel 44 does not intersect rotational axis 58 of aircraft 10. Rotor assemblies 60 and tail assemblies 46 are illustrated in non complementary configurations, however, regardless of the relative configuration of rotor assemblies 60 and tail assemblies 46, the horizontal components 68 of any thrust vectors 64 will not cause aircraft 10 to roll in any direction and will not cause aircraft 10 to rotate about rotational axis 58 as the passive brake configuration of tail assemblies 46 prevents such movement.

Referring to FIGS. 7A-7D in the drawings, the connections between propulsion assemblies 26 a, 26 b, 26 c, 26 d and airframe 12 will now be discussed. As illustrated, propulsion assemblies 26 a, 26 b, 26 c, 26 d are elements of the versatile propulsion system wherein, propulsion assemblies 26 a, 26 b, 26 c, 26 d are interchangeably attachable to airframe 12 as line replaceable units. Airframe 12 includes wings 14, 16 and pylons 18, 20. Pod assembly 50 is supported between pylons 18, 20 and is preferably rotatable and translatable relative to airframe 12. As illustrated, wings 14, 16 each have an M-wing design. Wing 14 has swept forward portions 14 c, 14 d and swept back portions 14 e, 14 f. Swept forward portion 14 c and swept back portion 14 e meet at leading apex 14 g. Swept forward portion 14 d and swept back portion 14 f meet at leading apex 14 h. Wing 16 has swept forward portions 16 c, 16 d and swept back portions 16 e, 16 f. Swept forward portion 16 c and swept back portion 16 e meet at leading apex 16 g. Swept forward portion 16 d and swept back portion 16 f meet at leading apex 16 h. Each of the swept forward portions 14 c, 14 d, 16 c, 16 d has a swept angle, which is depicted relative to swept forward portion 16 d as angle 70. Each of the swept back portions 14 e, 14 f, 16 e, 16 f has a swept angle, which is depicted relative to swept back portion 16 f as angle 72. Preferably, each of swept angles 70 are generally congruent with one another, each of swept angles 72 are generally congruent with one another and swept angles 70 are generally congruent with swept angles 72 such that airframe 12 is symmetric. In the illustrated embodiment, swept angles 70, 72 may preferably be between about 10 degrees and about 30 degrees, may more preferably be between about 15 degrees and about 25 degrees and may most preferably be about 20 degrees.

Airframe 12 includes four stanchions 74, only stanchions 74 a, 74 b positioned on the front of wing 14 being visible in FIG. 7A. Two similar stanchions are positioned aft of wing 16. In the illustrated embodiment, stanchion 74 a is located proximate leading apex 14 g and stanchion 74 b is located proximate leading apex 14 h. Similarly, the two stanchions of wing 16 are located proximate leading apex 16 g and proximate leading apex 16 h, respectively. Each of the stanchions 74 includes a flange 76 having a bolt pattern, only flanges 76 a, 76 b being visible in FIG. 7A. As best seen in FIG. 7C, stanchion 74 a includes an interface panel 78 a depicted with two power sockets 80 a, four data or communication sockets 82 a and two fluid sockets 84 a. As illustrated, sockets 80 a, 82 a, 84 a are substantially flush or integrated with panel 78 a. As should be apparent to those having ordinary skill in the art, each stanchion of the present disclosure will include a similar panel with similar sockets. In addition, even though a particular arrangement of sockets has been depicted and described, those having ordinary skill in the art should understand that the stanchions of the present disclosure could have other numbers of sockets in other arrangements.

Propulsion assemblies 26 a, 26 b, 26 c, 26 d each includes a flange 86, only flanges 86 c, 86 d being visible in FIG. 7A. Each flange 86 has a bolt pattern that matches the bolt pattern of flanges 76 such that propulsion assemblies 26 can be interchangeably bolted to any one of the stanchions 74 to create a mechanical connection therebetween. As best seen in FIG. 7D, propulsion assembly 26 a includes an interface panel 88 a depicted with two power cables 90 a, four data or communication cables 92 a and two fluid cables 94 a. Power cables 90 a are operable to couple with power sockets 80 a to established electrical connections between airframe 12 and propulsion assembly 26 a. For example, these connections enable electrical power from batteries 30 b of airframe 12 to be provided to components within propulsion assembly 26 a such as electronics node 40 a, an electric motor and/or other electrical components.

Communication cables 92 a are operable to couple with communication sockets 82 a to established data communication between airframe 12 and propulsion assembly 26 a. For example, these connections enable flight control system 40 to communicate with electronics node 40 a to provide command and control information to propulsion assembly 26 a and receive sensor and feedback information from propulsion assembly 26 a. Fluid cables 94 a are operable to couple with fluid sockets 84 a to established fluid communication between airframe 12 and propulsion assembly 26 a. For example, these connections enable liquid fuel from airframe 12 to be provided to a fuel tank and/or an internal combustion engine of propulsion assembly 26 a. Alternatively or additionally, these connections may enable hydraulic fluid from airframe 12 to provide power to hydraulic components within propulsion assembly 26 a. As should be apparent to those having ordinary skill in the art, each propulsion assembly of the present disclosure will include a similar panel with similar cables. In addition, even though a particular arrangement of cables has been depicted and described, those having ordinary skill in the art should understand that the propulsion assemblies of the present disclosure could have other numbers of cables in other arrangements that preferably mate with corresponding sockets of the stanchions of the present disclosure.

As illustrated, stanchions 74 provide standoff between propulsion assemblies 26 and wings 14, 16. By providing standoff between propulsion assemblies 26 and wings 14, 16, the aerodynamics of aircraft 10 are improved by effectively creating more wing surface to provide lift during various flight maneuvers. The M-wing design of wings 14, 16 and the attachment of propulsion assemblies 26 to stanchions 74 proximate forward apexes 14 g, 14 h, 16 g, 16 h enables to rotor assemblies 60 to be mounted in close proximity to forward apexes 14 g, 14 h, 16 g, 16 h and be tilted to generate variable thrust vectors 64. In the illustrated embodiment, each rotor assembly 60 has a tilting degree of freedom relative to the respective mast axis 56 enabling thrust vectors 64 to be resolved within a thrust vector cone about mast axis 56, as best seen in FIGS. 6A-6D. The maximum angle of the thrust vector may preferably be between about 10 degrees and about 30 degrees, may more preferably be between about 15 degrees and about 25 degrees and may most preferably be about 20 degrees. Importantly, the maximum thrust vector angle is preferably generally congruent with swept angles 70, 72 of wings 14, 16, thereby avoiding interference between rotor assemblies 60 and airframe 12.

Referring additionally to FIG. 8 in the drawings, a block diagram depicts an aircraft control system 100 operable for use with aircraft 10 of the present disclosure. In the illustrated embodiment, system 100 includes three primary computer based subsystems; namely, an airframe system 102, a passenger pod assembly system 104 and a remote system 106. As discussed herein, the aircraft of the present disclosure may be operated autonomously responsive to commands generated by flight control system 108 that preferably includes a non-transitory computer readable storage medium including a set of computer instructions executable by a processor. Flight control system 108 may be a triply redundant system implemented on one or more general-purpose computers, special purpose computers or other machines with memory and processing capability. For example, flight control system 108 may include one or more memory storage modules including, but is not limited to, internal storage memory such as random access memory, non-volatile memory such as read only memory, removable memory such as magnetic storage memory, optical storage, solid-state storage memory or other suitable memory storage entity. Flight control system 108 may be a microprocessor-based system operable to execute program code in the form of machine-executable instructions. In addition, flight control system 108 may be selectively connectable to other computer systems via a proprietary encrypted network, a public encrypted network, the Internet or other suitable communication network that may include both wired and wireless connections.

In the illustrated embodiment, flight control system 108 includes a command module 110 and a monitoring module 112. It is to be understood by those skilled in the art that these and other modules executed by flight control system 108 may be implemented in a variety of forms including hardware, software, firmware, special purpose processors and combinations thereof. Flight control system 108 receives input from a variety of sources including internal sources such as sensors 114, controllers 116 and propulsion assemblies 118-122, and external sources such as passenger pod assembly system 104, remote system 106 as well as global positioning system satellites or other location positioning systems and the like. For example, flight control system 108 may receive a flight plan including starting and ending locations for a mission from passenger pod assembly system 104 and/or remote system 106. Thereafter, flight control system 108 is operable to autonomously control all aspects of flight of an aircraft of the present disclosure.

For example, during the various operating modes of aircraft 10 including vertical takeoff and landing flight mode, hover flight mode, forward flight mode and transitions therebetween, command module 110 provides commands to controllers 116. These commands enable independent operation of each propulsion assembly 118-122 including tilting the rotor assemblies, adjusting the pitch of the proprotor blades, rotating, tilting and/or extending the tail assemblies and the like. Flight control system 108 receives feedback from controllers 116 and each propulsion assembly 118-122. This feedback is processes by monitoring module 112 that can supply correction data and other information to command module 110 and/or controllers 116. Sensors 114, such as positioning sensors, attitude sensors, speed sensors, environmental sensors, fuel sensors, temperature sensors, location sensors and the like also provide information to flight control system 108 to further enhance autonomous control capabilities.

Some or all of the autonomous control capability of flight control system 108 can be augmented or supplanted by remote flight control from, for example, remotes system 106 such as a transportation services provider. Remote system 106 may include one or computing systems that may be implemented on general-purpose computers, special purpose computers or other machines with memory and processing capability. For example, the computing systems may include one or more memory storage modules including, but is not limited to, internal storage memory such as random access memory, non-volatile memory such as read only memory, removable memory such as magnetic storage memory, optical storage memory, solid-state storage memory or other suitable memory storage entity. The computing systems may be microprocessor-based systems operable to execute program code in the form of machine-executable instructions. In addition, the computing systems may be connected to other computer systems via a proprietary encrypted network, a public encrypted network, the Internet or other suitable communication network that may include both wired and wireless connections. The communication network may be a local area network, a wide area network, the Internet, or any other type of network that couples a plurality of computers to enable various modes of communication via network messages using as suitable communication techniques, such as transmission control protocol/internet protocol, file transfer protocol, hypertext transfer protocol, internet protocol security protocol, point-to-point tunneling protocol, secure sockets layer protocol or other suitable protocol. Remote system 106 communicates with flight control system 108 via a communication link 124 that may include both wired and wireless connections.

Remote system 106 preferably includes one or more flight data display devices 126 configured to display information relating to one or more aircraft of the present disclosure. Display devices 126 may be configured in any suitable form, including, for example, liquid crystal displays, light emitting diode displays, cathode ray tube displays or any suitable type of display. Remote system 106 may also include audio output and input devices such as a microphone, speakers and/or an audio port allowing an operator to communicate with, for example, a pilot on board a pod assembly. The display device 126 may also serve as a remote input device 128 if a touch screen display implementation is used, however, other remote input devices, such as a keyboard or joystick, may alternatively be used to allow an operator to provide control commands to an aircraft being operated responsive to remote control.

Some or all of the autonomous and/or remote flight control of an aircraft of the present disclosure can be augmented or supplanted by onboard pilot flight control from an attached passenger pod assembly including system 104. Passenger pod assembly system 104 preferably includes a non-transitory computer readable storage medium including a set of computer instructions executable by a processor and may be implemented by a general-purpose computer, a special purpose computer or other machine with memory and processing capability. Passenger pod assembly system 104 may include one or more memory storage modules including, but is not limited to, internal storage memory such as random access memory, non-volatile memory such as read only memory, removable memory such as magnetic storage memory, optical storage memory, solid-state storage memory or other suitable memory storage entity. Passenger pod assembly system 104 may be a microprocessor-based system operable to execute program code in the form of machine-executable instructions. In addition, passenger pod assembly system 104 may be connectable to other computer systems via a proprietary encrypted network, a public encrypted network, the Internet or other suitable communication network that may include both wired and wireless connections. Passenger pod assembly system 104 communicates with flight control system 108 via a communication channel 130 that preferably includes a wired connection.

Passenger pod assembly system 104 preferably includes a cockpit display device 132 configured to display information to an onboard pilot. Cockpit display device 132 may be configured in any suitable form, including, for example, as one or more display screens such as liquid crystal displays, light emitting diode displays and the like or any other suitable display type including, for example, a display panel, a dashboard display, an augmented reality display or the like. Passenger pod assembly system 104 may also include audio output and input devices such as a microphone, speakers and/or an audio port allowing an onboard pilot to communicate with, for example, an operator of a remote system. Cockpit display device 132 may also serve as a pilot input device 134 if a touch screen display implementation is used, however, other user interface devices may alternatively be used to allow an onboard pilot to provide control commands to an aircraft being operated responsive to onboard pilot control including, for example, a control panel, mechanical control devices or other control devices. As should be apparent to those having ordinarily skill in the art, through the use of system 100, an aircraft of the present disclosure can be operated responsive to a flight control protocol including autonomous flight control, remote flight control or onboard pilot flight control and combinations thereof.

Referring next to FIGS. 9A-9B of the drawings, therein is depicted a rotor assembly for use on an aircraft 10 that is operable to generate a variable thrust output and a variable thrust vector at a constant rotational speed and that is generally designated 200. In the illustrated embodiment, rotor assembly 200 includes a mast 202 that is preferably rotated at a constant speed responsive to torque and rotational energy provided by the engine and drive system of the respective propulsion assembly. Mast 202 rotates about a mast axis 204. A ball joint 206 is positioned about mast 202 but does not rotate with mast 202. Instead, in the illustrated embodiment, ball joint 206 includes a flange 208 that is coupled to the airframe of aircraft 10 by bolting or other suitable technique. Ball joint 206 preferably has an outer spherical surface 210 that is operable to receive an inner spherical surface 211 of a tilt control assembly 212 thereon such that tilt control assembly 212 has a tilting degree of freedom relative to ball joint 206. Tilt control assembly 212 does not rotate with mast 202. Instead, a tilting plate 214 of tilt control assembly 212 is coupled to the airframe with a scissor mechanism 216 that includes a hinge 218 and a ball joint 220 that is received in a ball socket (not visible) of tilting plate 214. Two control rods 222 are coupled to tilting plate 214 such that control rods 222 are operable to push and pull tilting plate 214, thus actuating the tilting degree of freedom of tilt control assembly 212. Control rods 222 may be electrically, hydraulically and/or mechanically controlled responsive to flight control commands received from flight control system 108 via autonomous flight control, remote flight control, onboard pilot flight control or combinations thereof.

A rotor hub 224 is rotatably coupled to tilt control assembly 212 by a bearing assembly depicted as a ball bearing assembly 226, which provides for low friction relative rotation between rotor hub 224 and tilt control assembly 212. Rotor hub 224 is rotated by mast 202 via a rotational joint 228, such as a universal joint or a constant velocity joint, that is coupled to mast 202 by bolting or other suitable technique. Rotational joint 228 provides a torque path between mast 202 and rotor hub 224. Rotational joint 228 has a splined connector 230 that is received within a splined portion 231 of a drive arm assembly 232 of rotor hub 224. The splined mating surfaces allow rotor hub 224 to translate relative to rotational joint 228 and thus mast 202 during rotary operations. Rotor hub 224 rotates in a rotational plane about mast axis 204. The rotational plane may be normal to mast axis 204 when tilt control assembly 212 is not tilted relative to ball joint 206. In addition, the rotational plane may have an angle relative to mast axis 204 when tilt control assembly 212 is tilted relative to ball joint 206 as rotor hub 224 tilts with tilt control assembly 212 responsive to actuation of control rods 222. Rotor hub 224 including a plurality of spindle grips depicted as three spindle grips 234 a, 234 b, 234 c, in the illustrated embodiment. Spindle grips 234 a, 234 b, 234 c extend generally radially outwardly from the body of rotor hub 224. As best seen in the exploded section, spindle grip 234 b includes a spindle assembly 234 d having outboard and inboard bearings 234 e, 234 f onto which grip assembly 234 g is secured and operable to rotate thereabout. Spindle grips 234 a, 234 c preferably have similar construction and operation such that spindle grips 234 a, 234 b, 234 c are operable to rotate about respective pitch change axes 236 a, 236 b, 236 c. In the illustrated embodiment, each spindle grip 234 a, 234 b, 234 c includes a respective pinion gear 238 a, 238 b, 238 c. A rotor blade (see for example FIGS. 6A-6D) is coupled to each of the spindle grips 234 a, 234 b, 234 c at respective devises 240 a, 240 b, 240 c by bolting, pinning or other suitable technique. The pitch of the rotor blades is controlled responsive to rotation of the spindle grips 234 a, 234 b, 234 c about the respective pitch change axes 236 a, 236 b, 236 c.

Rotor assembly 200 includes a collective pitch control mechanism 242. In the illustrated embodiment, collective pitch control mechanism 242 includes a spider assembly 244 having a plurality of arms extending therefrom depicted as three arms 246 a, 246 b, 246 c. Each arm 246 a, 246 b, 246 c includes a rack gear 248 a, 248 b, 248 c that is operable to mate with a respective pinion gear 238 a, 238 b, 238 c such that translation of spider assembly 244 relative to rotor hub 224 responsive to operation of one or more actuators 250 rotates each spindle grip 234 a, 234 b, 234 c about the respective pitch change axis 236 a, 236 b, 236 c to collectively adjust the pitch of the rotor blades, thereby generating the variable thrust output at a constant rotational speed. In addition, actuation of tilt control assembly 212 changes the rotational plane of rotor hub 224 relative to mast axis 204, thereby generating the variable thrust vector. As the control rods 222 are operable to tilt tilting plate 214 in any direction relative to ball joint 206, the rotational plane of rotor hub 224 may be tilted in any direction relative to mast axis 204 thus enabling resolution of the thrust vector within a thrust vector cone relative to mast axis 204. In some embodiments, the thrust vector cone may have a maximum angle relative to mast axis 204 of between about ten degrees and about thirty degrees. In other embodiments, the thrust vector cone may have a maximum angle relative to mast axis 204 of between about fifteen degrees and about twenty-five degrees. In additional embodiments, the thrust vector cone may have a maximum angle relative to mast axis 204 of about twenty degrees. It is noted that when rotor assemblies 200 are used in the versatile propulsion system of aircraft 10, each rotor assembly 200 is positioned proximate a leading apex in the M-wing design thus enabling the disclosed resolution of the thrust vector within the thrust vector cone without interference between the rotor blades and the airframe.

Referring next to FIGS. 10A-10B of the drawings, therein is depicted a rotor assembly for use on an aircraft 10 that is operable to generate a variable thrust output and a variable thrust vector at a constant rotational speed and that is generally designated 300. In the illustrated embodiment, rotor assembly 300 includes a mast 302 that is preferably rotated at a constant speed responsive to torque and rotational energy provided by the engine and drive system of the respective propulsion assembly. Mast 302 rotates about a mast axis 304. A ball joint 306 is positioned about mast 302 but does not rotate with mast 302. Instead, in the illustrated embodiment, ball joint 306 includes a flange 308 that is coupled to the airframe of aircraft 10 by bolting or other suitable technique. Ball joint 306 preferably has an outer spherical surface 310 that is operable to receive an inner spherical surface 311 of a tilt control assembly 312 thereon such that tilt control assembly 312 has a tilting degree of freedom relative to ball joint 306. Tilt control assembly 312 does not rotate with mast 302. Instead, a tilting plate 314 of tilt control assembly 312 is coupled to the airframe with a scissor mechanism 316 that includes a hinge 318 and a ball joint 320 that is received in a ball socket (not visible) of tilting plate 314. Two control rods 322 are coupled to tilting plate 314 such that control rods 322 are operable to push and pull tilting plate 314, thus actuating the tilting degree of freedom of tilt control assembly 312. Control rods 322 may be electrically, hydraulically and/or mechanically controlled responsive to flight control commands received from flight control system 108 via autonomous flight control, remote flight control, onboard pilot flight control or combinations thereof.

A rotor hub 324 is rotatably coupled to tilt control assembly 312 by a bearing assembly depicted as a ball bearing assembly 326, which provides for low friction relative rotation between rotor hub 324 and tilt control assembly 312. Rotor hub 324 is rotated by mast 302 via a rotational joint 328, such as a universal joint or a constant velocity joint, that is coupled to mast 302 by bolting or other suitable technique. Rotational joint 328 provides a torque path between mast 302 and rotor hub 324. Rotational joint 328 has a splined connector 330 that is received within a splined portion 331 of a drive arm assembly 332 of rotor hub 324. The splined mating surfaces allow rotor hub 324 to translate relative to rotational joint 328 and thus mast 302 during rotary operations. Rotor hub 324 rotates in a rotational plane about mast axis 304. The rotational plane may be normal to mast axis 304 when tilt control assembly 312 is not tilted relative to ball joint 306. In addition, the rotational plane may have an angle relative to mast axis 304 when tilt control assembly 312 is tilted relative to ball joint 306 as rotor hub 324 tilts with tilt control assembly 312 responsive to actuation of control rods 322. Rotor hub 324 including a plurality of spindle grips depicted as three spindle grips 334 a, 334 b, 334 c, in the illustrated embodiment. Spindle grips 334 a, 334 b, 334 c extend generally radially outwardly from the body of rotor hub 324. As best seen in the exploded section, spindle grip 334 b includes a spindle assembly 334 d having outboard and inboard bearings 334 e, 334 f onto which grip assembly 334 g is secured and operable to rotate thereabout. Spindle grips 334 a, 334 c preferably have similar construction and operation such that spindle grips 334 a, 334 b, 334 c are operable to rotate about respective pitch change axes 336 a, 336 b, 336 c. As illustrated, each spindle grip 334 a, 334 b, 334 c includes a respective pinion gear 338 a, 338 b, 338 c. A rotor blade (see for example FIGS. 6A-6D) is coupled to each of the spindle grips 334 a, 334 b, 334 c at respective devises 340 a, 340 b, 340 c by bolting, pinning or other suitable technique. The pitch of the rotor blades is controlled responsive to rotation of the spindle grips 334 a, 334 b, 334 c about the respective pitch change axes 336 a, 336 b, 336 c.

Rotor assembly 300 includes a collective pitch control mechanism 342. In the illustrated embodiment, collective pitch control mechanism 342 includes a ring assembly 344 having an input gear 346 and an output gear 348 that is operable to mate with each of the pinion gears 338 a, 338 b, 338 c such that rotation of ring assembly 344 relative to rotor hub 324 responsive to operation of one or more actuators 350 rotates each spindle grip 334 a, 334 b, 334 c about the respective pitch change axis 336 a, 336 b, 336 c to collectively adjust the pitch of the rotor blades, thereby generating the variable thrust output at a constant rotational speed. In addition, actuation of tilt control assembly 312 changes the rotational plane of rotor hub 324 relative to mast axis 304, thereby generating the variable thrust vector. As the control rods 322 are operable to tilt tilting plate 314 in any direction relative to ball joint 306, the rotational plane of rotor hub 324 may be tilted in any direction relative to mast axis 304 thus enabling resolution of the thrust vector within a thrust vector cone relative to mast axis 304. In some embodiments, the thrust vector cone may have a maximum angle relative to mast axis 304 of between about ten degrees and about thirty degrees. In other embodiments, the thrust vector cone may have a maximum angle relative to mast axis 304 of between about fifteen degrees and about twenty-five degrees. In additional embodiments, the thrust vector cone may have a maximum angle relative to mast axis 304 of about twenty degrees. It is noted that when rotor assemblies 300 are used in the versatile propulsion system of aircraft 10, each of rotor assembly 300 is positioned proximate a leading apex in the M-wing design thus enabling the thrust vector to be resolved within the thrust vector cone without interference between the rotor blades and the airframe.

The foregoing description of embodiments of the disclosure has been presented for purposes of illustration and description. It is not intended to be exhaustive or to limit the disclosure to the precise form disclosed, and modifications and variations are possible in light of the above teachings or may be acquired from practice of the disclosure. The embodiments were chosen and described in order to explain the principals of the disclosure and its practical application to enable one skilled in the art to utilize the disclosure in various embodiments and with various modifications as are suited to the particular use contemplated. Other substitutions, modifications, changes and omissions may be made in the design, operating conditions and arrangement of the embodiments without departing from the scope of the present disclosure. Such modifications and combinations of the illustrative embodiments as well as other embodiments will be apparent to persons skilled in the art upon reference to the description. It is, therefore, intended that the appended claims encompass any such modifications or embodiments. 

What is claimed is:
 1. A rotor assembly for an aircraft operable to generate a variable thrust output at a constant rotational speed, the rotor assembly comprising: a mast rotatable at the constant speed about a mast axis; a rotor hub coupled to and rotatable with the mast, the rotor hub including a plurality of spindle grips extending generally radially outwardly, each of the spindle grips operable to rotate about a pitch change axis; a plurality of rotor blades each coupled to one of the spindle grips and rotatable about the respective pitch change axis; and a collective pitch control mechanism coupled to and rotatable with the rotor hub, the collective pitch control mechanism operably associated with each spindle grip; wherein, actuation of the collective pitch control mechanism rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades, thereby generating the variable thrust output.
 2. The rotor assembly as recited in claim 1 wherein the collective pitch control mechanism further comprises a spider assembly having a plurality of arms each of which mate with one of the spindle grips such that translation of the spider assembly relative to the rotor hub rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades.
 3. The rotor assembly as recited in claim 2 wherein the collective pitch control mechanism further comprises at least one actuator operable to translate the spider assembly relative to the rotor hub.
 4. The rotor assembly as recited in claim 2 wherein each spindle grip further comprises a pinion gear, wherein each arm further comprises a rack gear and wherein each rack gear is operable to mate with one of the pinion gears.
 5. The rotor assembly as recited in claim 1 wherein the collective pitch control mechanism further comprises a ring assembly that mates with each of the spindle grips such that rotation of the ring assembly relative to the rotor hub rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades.
 6. The rotor assembly as recited in claim 5 wherein the ring assembly further comprises an input gear and wherein the collective pitch control mechanism further comprises at least one actuator having a drive gear operable to mate with the input gear and rotate the ring assembly relative to the rotor hub.
 7. The rotor assembly as recited in claim 5 wherein each spindle grip further comprises a pinion gear, wherein the ring assembly further comprises an output gear and wherein the output gear is operable to mate with each of the pinion gears.
 8. The rotor assembly as recited in claim 1 further comprising a joint assembly providing a torque path from the mast to the rotor hub.
 9. The rotor assembly as recited in claim 8 wherein the joint assembly further comprises a constant velocity joint assembly providing the torque path from the mast to the rotor hub.
 10. The rotor assembly as recited in claim 9 wherein the rotor hub further comprises a drive arm assembly coupled to the constant velocity joint assembly.
 11. The rotor assembly as recited in claim 10 wherein the drive arm assembly is operable to rotate with and translate relative to the constant velocity joint assembly during rotary operations.
 12. The rotor assembly as recited in claim 1 further comprising a ball joint positioned about and non rotatable with the mast and a tilt control assembly positioned on and having a tilting degree of freedom relative to the ball joint, the tilt control assembly non rotatable with the mast; and wherein, the rotor hub is rotatably coupled to and tiltable with the tilt control assembly such that actuation of the tilt control assembly changes the rotational plane of the rotor hub relative to the mast axis.
 13. An aircraft comprising: an airframe; a propulsion system attached to the airframe, the propulsion system including a plurality of propulsion assemblies each having a rotor assembly; and a flight control system operable to independently control each of the propulsion assemblies; the rotor assemblies each comprising: a mast rotatable relative to the airframe about a mast axis; a rotor hub coupled to and rotatable with the mast, the rotor hub including a plurality of spindle grips extending generally radially outwardly, each of the spindle grips operable to rotate about a pitch change axis; a plurality of rotor blades each coupled to one of the spindle grips and rotatable about the respective pitch change axis; and a collective pitch control mechanism coupled to and rotatable with the rotor hub, the collective pitch control mechanism operably associated with each spindle grip; wherein, actuation of the collective pitch control mechanism rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades, thereby generating a variable thrust output.
 14. The aircraft as recited in claim 13 wherein each of the collective pitch control mechanisms further comprises a spider assembly having a plurality of arms each of which mate with one of the spindle grips such that translation of the spider assembly relative to the rotor hub rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades.
 15. The aircraft as recited in claim 14 wherein each of the collective pitch control mechanisms further comprises at least one actuator operable to translate the spider assembly relative to the rotor hub; wherein, each spindle grip further comprises a pinion gear; wherein, each arm further comprises a rack gear; and wherein, each rack gear is operable to mate with one of the pinion gears.
 16. The aircraft as recited in claim 13 wherein each of the collective pitch control mechanisms further comprises a ring assembly that mates with each of the spindle grips such that rotation of the ring assembly relative to the rotor hub rotates each spindle grip about the respective pitch change axis to collectively control the pitch of the rotor blades.
 17. The aircraft as recited in claim 16 wherein each of the ring assemblies further comprises an input gear and an output gear; wherein, each of the collective pitch control mechanisms further comprises at least one actuator having a drive gear operable to mate with the input gear and rotate the ring assembly relative to the rotor hub; wherein, each spindle grip further comprises a pinion gear; and wherein, the output gear of the ring assembly is operable to mate with each of the pinion gears.
 18. The aircraft as recited in claim 13 wherein the airframe further comprises first and second wings each having an M-wing design with two leading apexes, each of the rotor assemblies positioned proximate one of the leading apexes.
 19. The aircraft as recited in claim 13 further comprising a passenger pod assembly rotatably attachable to the airframe such that the passenger pod assembly remains in a generally horizontal attitude in a vertical takeoff and landing fight mode, a forward flight mode and transitions therebetween.
 20. The aircraft as recited in claim 13 wherein the flight control system commands operation of the propulsion assemblies responsive to at least one of onboard pilot flight control, remote flight control, autonomous flight control and combinations thereof. 